XFOIL Version 6.94 Calculated polar for: N-9 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4652 0.00747 0.00212 -0.0658 0.7150 1.0000 0.500 0.5178 0.00777 0.00225 -0.0649 0.6826 1.0000 1.000 0.5703 0.00805 0.00240 -0.0640 0.6479 1.0000 1.500 0.6225 0.00833 0.00255 -0.0631 0.6067 1.0000 2.000 0.6737 0.00866 0.00272 -0.0620 0.5444 1.0000 2.500 0.7220 0.00933 0.00299 -0.0606 0.4516 1.0000 3.000 0.7711 0.01011 0.00348 -0.0595 0.3883 1.0000 3.500 0.8212 0.01083 0.00402 -0.0587 0.3411 1.0000 4.000 0.8710 0.01157 0.00461 -0.0578 0.2849 1.0000 4.500 0.9143 0.01313 0.00550 -0.0563 0.1523 1.0000 5.000 0.9597 0.01447 0.00671 -0.0549 0.1292 1.0000 5.500 1.0044 0.01584 0.00808 -0.0534 0.1166 1.0000 6.000 1.0491 0.01718 0.00946 -0.0520 0.1037 1.0000 6.500 1.0951 0.01824 0.01065 -0.0509 0.0919 1.0000 7.000 1.1435 0.01888 0.01148 -0.0501 0.0796 1.0000 7.500 1.1963 0.01900 0.01177 -0.0501 0.0559 1.0000 8.000 1.2307 0.02128 0.01385 -0.0473 0.0260 1.0000 8.500 1.2603 0.02376 0.01652 -0.0438 0.0209 1.0000 9.000 1.2774 0.02710 0.02004 -0.0390 0.0180 1.0000 9.500 1.2947 0.02994 0.02319 -0.0342 0.0168 1.0000 10.000 1.3065 0.03334 0.02689 -0.0293 0.0157 1.0000 10.500 1.3143 0.03720 0.03100 -0.0254 0.0146 1.0000 11.000 1.3139 0.04251 0.03659 -0.0219 0.0138 1.0000 11.500 1.3023 0.04975 0.04431 -0.0191 0.0134 1.0000 12.000 1.2812 0.05826 0.05331 -0.0185 0.0133 1.0000 12.500 1.2568 0.06780 0.06331 -0.0206 0.0133 1.0000 13.000 1.2247 0.07958 0.07550 -0.0251 0.0133 1.0000 13.500 1.1914 0.09296 0.08925 -0.0320 0.0134 1.0000 14.000 1.1626 0.10736 0.10397 -0.0407 0.0135 1.0000 14.500 1.1313 0.12409 0.12101 -0.0513 0.0137 1.0000